This invention relates to aircraft navigation systems and, more particularly, to an apparatus and method for aircraft navigation that utilizes a blended architecture consisting of a global positioning system (GPS) and micro-electromechanical sensors (MEMS) for the primary navigation system and a laser gyroscope system for the secondary navigation system.
An aircraft navigation system is the source of data for many critical avionics functions, such as the primary flight control system, the flight deck displays, and guidance, control, and stabilization systems, including automatic landing systems. The navigation system measures a variety of parameters defining the state of the aircraft, such as attitude, heading, angular rates, acceleration, track angle, flight path angle, ground speed and position, and provides the data to the avionics systems for display and for use in the control of the aircraft""s flight.
Commercial aircraft have generally relied upon inertial navigation. Inertial navigation requires that the navigation system be initialized at a starting position and provide autonomous and continuous measurements based on that reference. As such, inertial navigation systems are particularly useful for over-water navigation where it is more difficult to obtain ground references for the measurements. Most inertial navigation systems, however, are expensive and are subject to increases in position error, commonly called xe2x80x9cdrift,xe2x80x9d over time.
More particularly, most modem commercial aircraft are equipped with traditional or, less commonly, fault-tolerant Air Data Inertial Reference Units (ADIRU) to perform stand-alone inertial navigation and provide the necessary air data to the avionics systems. To ensure that navigational data is continuously provided during a flight, aircraft generally have more than one, typically three, traditional ADIRUs operating in parallel in a redundant arrangement, called a triplex configuration. Such traditional ADIRUs are used in the majority of large commercial aircraft. In this regard, those aircraft having Category 3B automatic landing capability require three ADIRUs, while the other navigational requirements can be minimally met with two ADIRUs. Alternatively, the most modern commercial aircraft may have a single, fault-tolerant ADIRU configuration that is constructed to be equivalent to three separate traditional ADIRUs. The single fault-tolerant ADIRU is constructed such that if any one component fails, the ADIRU remains operational, since the same ADIRU includes redundant components. In fact, the fault-tolerant ADIRU is generally constructed to remain operational even if any two components fail.
Each traditional ADIRU has three navigation-grade ring laser gyroscopes and three accelerometers. Therefore, a triplex configuration of the traditional ADIRUs has a total of nine navigation-grade ring laser gyroscopes and nine accelerometers. Each fault-tolerant ADIRU, on the other hand, has six navigation-grade ring laser gyroscopes and six accelerometers. The navigation-grade ring laser gyroscopes and accelerometers provide inertial navigation for the aircraft with a low amount of drift, typically less than 0.01 degree/hour, but they are expensive. All ADIRUs, traditional and fault-tolerant, require that the inertial measurements be obtained with great precision and that subsequent processing of those measurements maintain that precision. Thus, ADIRU processors generally have a complex and proprietary sensor interface to provide the precise timing, measurements and specialized features that are necessary. In addition, the processors and interfaces must generally be manufactured or provided by the same company that provided the sensors to ensure compatibility among the components of the ADIRU. ADIRUs are available from various vendors including the Honeywell HG2050 (traditional ADIRU) and HG2060 (fault-tolerant ADIRU) and the Litton (Northrop Grumman) LTN-101 (traditional ADIRU).
An aircraft equipped with a fault-tolerant ADIRU also carries a Secondary Attitude Air data Reference Unit (SAARU), which is a backup to the ADIRU in the rare event that the ADIRU malfunctions. This architecture is called a Fault-Tolerant Air Data Inertial Reference System (FT-ADIRS). The components of the SAARU are intentionally dissimilar to the ADIRU to preclude common failures in both units. That is, the SAARU generally will not include ring laser gyroscopes if the ADIRU includes ring laser gyroscopes. The SAARU may have four attitude-grade fiber optic gyroscopes. The fiber optic gyroscopes provide the necessary dissimilar design, but suffer from a higher amount of drift than the laser gyroscopes of the ADIRU with the drift generally being several degrees/hour. In addition, fiber optic gyroscopes are also costly. Like the laser gyroscopes of the ADIRU, the fiber optic gyroscope configuration of the SAARU requires its own processors, power supplies, input/output modules and proprietary interface to process signals, which also increases the cost of the SAARU. Furthermore, the SAARU is not fault-tolerant, so it must be fully functional before aircraft operation to ensure the availability of this backup unit.
By way of example, FIG. 1 depicts a conventional FT-ADIRS architecture, located in the forward electrical/electronics bay of a Boeing 777 aircraft, with a single, fault-tolerant ADIRU 12. The ADIRU 12 has six ring laser gyroscopes and six accelerometers designated generally as 14 and four processors 16 (P1-P4) to process the signals from the gyroscopes and accelerometers 14. The processors 16 communicate with other avionics systems via the flight control buses, which generally include a right flight control bus 26, a center flight control bus 28 and a left flight control bus 30. The three flight control buses conform to ARINC standard 629 and are high-speed, two-way data buses that are shared by all subscribing ARINC 629 input/output (I/O) terminals attached to them. The I/O modules in the fault-tolerant ADIRU serve to receive air data sensor inputs and to transmit the entire suite of inertial and air data state signals to user avionics systems, such as flight instruments, flight management, automatic pilot controls and primary flight controls. Since signals may be transmitted both to and from the ADIRU 12 via the right 26 and left 30 flight control bus, transceivers 20, 24 are generally disposed between the respective input/output modules 18 and the right 26 and left 30 flight control bus. In contrast, since signals are generally received by the ADIRU 12 from the center flight control bus 28, a receiver 22 is generally disposed between the respective input/output modules 18 and the center flight control bus 28. To provide the desired redundancy, the ADIRU 12 generally includes multiple redundant processors 16, at least two input/output modules 18 associated with each flight control bus 26, 28, 30 and at least two transceivers or receivers 20, 22, 24 in communication with the respective flight control bus.
FIG. 1 also shows the backup SAARU 32 with components that are dissimilar to the components of the ADIRU 12. The SAARU 32 has four fiber optic gyroscopes 34 and two processors 36 (PY, PZ) to process the signals from the gyroscopes 34. The SAARU 32 also includes multiple input/output modules 38, one of which is associated with each flight control bus 26, 28, 30. The respective input/output modules 38 are, in turn, connected to the right 26 and left 30 flight control bus by respective receivers 40, 44 and to the center flight control bus 28 by a transceiver 42.
The flight control buses are also connected to the left Aircraft Information Management System (AIMS) 52 and the right AIMS 50. The left AIMS 52 is generally connected to the left flight control bus 30 by a transceiver 64 and to the center 28 and right 26 flight control bus by a receiver 60, 62. Conversely, the right AIMS 50 is generally connected to the right flight control bus 26 by a transceiver 58 and to the left 30 and center 28 flight control bus by a receiver 54, 56. As FIG. 1 illustrates, the left AIMS 52 is connected to and receives data from the Total Air Temperature (TAT) probe 49 and the left Angle Of Attack (AOA) vane 48. The right AIMS 50 is also connected to and receives information from the TAT probe 49 and the right AOA vane 46. Thus, the right and left AIMS 50, 52 provide an interface by converting the analog outputs of the AOA vane 48 and TAT probe 49 to digital signals for transmission to the ADIRU 12 and SAARU 32 via the flight control buses.
In addition, the flight control buses are connected to six primary air data modules (ADM). These modules measure the total and static pressure associated with the air surrounding the aircraft and provide signals to the ADIRU 12 and SAARU 32 via the flight control buses. ADM 68 measures total pressure from the center pitot probe 67 with ADM 68 being connected directly to the center flight control bus 28 by a transceiver 84. As known to those skilled in the art, a pitot probe generally measures total pressure. ADM 70 similarly measures total pressure from the left pitot probe 69 and is connected directly to the left flight control bus 30 by a transceiver 86. Likewise, ADM 66 measures total pressure from the right pitot probe 65 and connects to the right flight control bus 26 by a transceiver 82. ADM 72, 74, 76 and STBY ADM 78 also measure the static pressure via static ports 94. ADM 72 is, in turn, connected to the left flight control bus 30 by transceiver 88, ADM 74 is connected to the right flight control bus 26 by transceiver 90, and ADM 76 is connected to the center flight control bus 28 by transceiver 92.
In addition to the flight control bus, standby display buses are also depicted in FIG. 1 as dashed lines that provide information to the standby displays 100 in the rare event the primary displays or the ARINC 629 flight control buses malfunction. The standby display buses conform to ARINC standard 429 and connect the SAARU 32 and STBY ADMs 78 and 80 to the standby displays 100. The ARINC 429 bus provides a means dissimilar to the ARINC 629 bus for transmitting critical attitude data from the SAARU 32, through the 429 module 98, to the standby attitude display that also is dissimilar to the primary displays.
As a further consideration, both the ADIRU and SAARU are capable of measuring the aircraft""s pitch rate. The ADIRU and SAARU pitch rate measurements, however, may not be suitable for the primary flight control system because the ADIRU and SAARU are located in the forward electrical/electronics bay of the aircraft, which is also subject to the aircraft""s longitudinal bending modes. Because of their location, the ADIRU and SAARU measurements of the aircraft pitch rate contain an unacceptably high level of signal representing the motion associated with the aircraft""s longitudinal bending modes, creating unsuitable pitch rate measurements. Thus, to provide the primary flight control system with suitable pitch rate measurements that are relatively free of longitudinal bending motion, four additional pitch rate sensors are installed at the wing main spar which experiences much less structural bending motion.
As will be apparent, the FT-ADIRS suite, containing the fault-tolerant ADIRU, SAARU, six primary air data modules, two standby air data modules, and four additional pitch rate sensors is expensive, often costing more than $200,000. As a result of the significant cost of a conventional commercial aircraft inertial navigation system, there exists a need in the aircraft industry for a much less expensive, but equally or more precise and redundant, navigation system. In addition, the industry would benefit greatly from a navigation system created by a relatively simple combination of components that does not require dedicated processors that communicate via a proprietary interface.
In accordance with this invention, an apparatus and method for navigation of an aircraft are provided that utilize a blended architecture having a global positioning system (GPS) and micro-electromechanical sensors (MEMS) for the redundant, primary navigation system and a laser gyroscope system for the secondary navigation system. The structure of the navigation system of the present invention provides a relatively low-cost navigation system because GPS and MEMS components are less expensive than the laser gyroscopes and accelerometers that are utilized in the conventional inertial navigation systems, such as the FT-ADIRS, utilized by commercial aircraft. Although the present invention utilizes laser gyroscopes in the secondary navigation system, fewer laser gyroscopes are necessary for a secondary navigation- system than in conventional navigation systems. The navigation system of the present invention may also have a lower cost because the primary navigation system does not require a proprietary sensor interface and may share processing resources with other avionics systems, unlike conventional systems that have a proprietary sensor interface and require dedicated processors. In addition, the primary navigation system of the present invention need not be centralized and may be distributed throughout the aircraft at the optimal locations for the respective components, such that the primary navigation system may measure angular rates, such as the pitch rate, and accelerations from locations in the aircraft that are optimal for the type of measurement at issue. This eliminates the cost of extra components that are necessary to measure the pitch rate in a conventional navigation system. Overall, the navigation system of the present invention is significantly less expensive and easier to maintain than the conventional navigation systems utilized by commercial aircraft because the components cost less and supporting systems may be shared with other avionics systems.
According to the present invention, the apparatus for navigation of an aircraft comprises a primary navigation system and a secondary navigation system. The primary system may have redundant global positioning system (GPS) receivers to provide signals indicative of a position of the aircraft and redundant micro-electromechanical sensors (MEMS) to perform inertial referencing of the aircraft. The GPS receivers may also determine the heading of the aircraft. The MEMS may include MEMS rate sensors and MEMS accelerometers. The secondary navigation system may be a single laser gyroscope system, typically including three ring laser gyroscopes and three navigation-grade accelerometers, that continuously performs inertial referencing of the aircraft and produces purely inertial navigation signals. The output of the primary navigation system is utilized to provide the necessary data to all of the avionics systems. However, the signals provided by the secondary navigation system may be utilized in the rare event the redundant primary navigation system completely malfunctions. By typically including only a single traditional ADIRU and instead relying primarily upon redundant GPS and the MEMS sensors, the apparatus of the present invention is more economical.
The apparatus of the present invention may also include a general-purpose computer that supports the primary navigation system in addition to other avionics systems. The general-purpose computer may have redundant processors. The apparatus of this embodiment also significantly reduces the price of an aircraft navigation system because a dedicated computer is not required to process the signals from the primary navigation system, unlike the conventional inertial navigation systems for commercial aircraft that require dedicated processors to provide the necessary precision. Moreover, since the primary navigation system does not rely upon the precision offered by ring laser gyroscopes, the apparatus of the present invention also need not include the complex and proprietary interface included within conventional navigation systems for commercial aircraft.
The MEMS rate sensors and MEMS accelerometers may be distributed throughout the aircraft. The MEMS rate sensors and MEMS accelerometers may be distributed in clusters of three MEMS rate sensors and three MEMS accelerometers. Since each MEMS rate sensor may provide an output based upon motion of the aircraft in a respective direction, the clusters of three MEMS rate sensors may be mounted such that the directions in which the respective MEMS rate sensors sense motion of the aircraft are mutually orthogonal. Likewise, since each MEMS accelerometer may provide an output based upon motion of the aircraft in a respective direction, the clusters of three MEMS accelerometers may be mounted such that the directions in which the respective MEMS accelerometers sense motion of the aircraft are mutually orthogonal. In a further embodiment of the present invention, the three MEMS rate sensors of a cluster may be mounted such that each senses motion of the aircraft along a respective one of the three, mutually orthogonal principle axes of the aircraft. In the same way, the three MEMS accelerometers of a cluster may be mounted such that each senses motion of the aircraft along a respective one of the three, mutually orthogonal principle axes of the aircraft. This distributed arrangement of the MEMS sensors is advantageous since the clusters may be placed in locations where the most suitable measurements will be obtained for use by the avionics systems, for example the primary flight control system. This is a significant improvement over the conventional approach of housing the FT-ADIRS entirely within the forward electrical/electronics bay of the aircraft and then installing additional sensors at the main wing spar to measure the pitch rate of the aircraft.
In an embodiment of the method of the present invention, a primary navigation system having redundant GPS and MEMS sensors and a secondary navigation system having a single laser gyroscope system are initially provided. During flight, the GPS may then provide signals indicative of, among other things, the position, groundspeed and the heading of the aircraft and the MEMS sensors may provide signals indicative of inertial motion of the aircraft. According to the method of the present invention, the aircraft is navigated, guided and controlled based upon the signals provided by the GPS and MEMS sensors so long as the GPS and MEMS sensors are operational. The signals provided by the redundant GPS and MEMS sensors may be consolidated and blended in a suitable filter to derive a single suite of high-integrity navigation data for use in the avionics systems. In the rare event the redundant primary navigation system completely malfunctions, the aircraft may be navigated, guided and controlled based upon signals from the laser gyroscope system indicative of the inertial motion of the aircraft.
Thus, the apparatus and method of the present invention provides the industry with a less expensive, but equally or more precise and redundant, navigation system relative to the conventional navigation systems utilized by commercial aircraft. The industry will benefit greatly from the primary navigation system of the present invention created by a combination of low-cost GPS and MEMS components that do not require dedicated processors or a proprietary interface. Although a conventional gyroscope configuration is utilized for the secondary navigation system, the overall design of the navigation system of the present invention nevertheless decreases the cost because fewer expensive gyroscopes, accelerometers and dedicated processors are necessary. Thus, the industry need for a navigation system with components that may be distributed throughout the aircraft and that may share computing resources with other avionics systems to process signals and provide data to the avionics systems is met by the present invention.